1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with film cooling slots.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an industrial gas turbine (IGT) engine, includes a turbine with multiple rows or stages or stator vanes that guide a high temperature gas flow through adjacent rotors of rotor blades to produce mechanical power and drive a bypass fan, in the case of an aero engine, or an electric generator, in the case of an IGT. In both cases, the turbine is also used to drive the compressor.
It is well known in the art of gas turbine engine design that the efficiency of the engine can be increased by passing a higher gas flow temperature through the turbine. However, the turbine inlet temperature is limited by the material properties of the turbine, especially for the first stage airfoils since these are exposed to the highest temperature gas flow. As the gas flow passes through the various stages of the turbine, the temperature decreases as the energy is extracted by the rotor blades.
Another method of increases the turbine inlet temperature is to provide more effective cooling of the airfoils. Complex internal and external cooling circuits or designs have been proposed using a combination of internal convection and impingement cooling along with external film cooling to transfer heat away from the metal and form a layer of protective air to limit thermal heat transfer to the metal airfoil surface. However, since the pressurized air used for the airfoil cooling is bled off from the compressor, this bleed off air decreases the efficiency of the engine because the work required to compress the air is not used for power production. It is therefore wasted energy as far as producing useful work in the turbine.
Shaped diffusion film cooling holes are normally used for the cooling of a turbine blade pressure side wall. The use of axial oriented film cooling holes on the pressure side surface of the airfoil is mainly for an injection of cooling air inline with the main stream gas flow which is accelerated into multiple directions as represented by the various arrows in FIG. 1.
However, at the airfoil pressure side surface two thirds of the way downstream from the leading edge region, the hot gas secondary flow migrates in the multiple directions, depending on the pressure gradient and also moving in an axial direction. Due to pressure gradient across the blade tip, the upper blade span hot gas flow migrates toward the blade tip section. Due to the nature of turbine expansion, the middle portion flow in the axial direction. Due to the hot gas passage channel pressure gradient, the lower span hot gas flow migrates toward the blade platform. An axial oriented shaped film cooling hole is used in that region of the blade thus becomes undesirable. FIG. 1 shows the secondary hot gas flow phenomena on the blade pressure side surface.